Axial flow compressor



Dec. 26, 1961 Filed June 9, 1958 4 Sheets-Sheet 1 INVENTORS Dec. 26,1961 J. N. BARNEY ETAL 3,014,640

AXIAL FLOW COMPRESSOR Filed June 9, 1958 4 Sheets-Sheet 2 INVENTORS 5ATTORNEY Dec. 26, 196 1 J. N. BARNEY ETAL 3,014,640

AXIAL FLOW COMPRESSOR Filed June 9, 1958 4 Sheets-Sheet 3 co nsereo meFzaw EFFIC7NCY ALTITUDE L033 COEFFICIENT REYNOLDS NUMBER ATTORNEY Dec.26, 1961 J. N. BARNEY ETAL 3,014,640 AXIAL FLOW COMPRESSOR Filed June 9,1958 4 Sheets-Sheet 4 M4! STREAM FLOW a 31 STREAM 10W 7? w H x %g 1 5 X;

v A INVENTORS 222$2 Our invention relates to turbocompressors, and isdirected to improving the performance of such machines, particularlyaxial-flow compressors.

The invention relates particularly to improving the characteristics ofcompressors by controlling boundary layer effects to improve flowcharacteristics and efiiciency.

The nature of the invention and its advantages will be clear to thoseskilled in the art from the succeeding detailed description of thepreferred embodiment of the invention and the accompanying drawingsthereof.

FIGURE 1 is a partial sectional view of an axial-flow compressor takenon a. plane containing the axis of the compressor.

FIGURE 2 is an axonometric view of a rotor blade.

FIGURE 3 is an elevation View of the base portion of a rotor blade.

v FIGURE 4 is a cross-sectional view of a rotor blade taken on the planeindicated by the line 4-4 in FIG- URE 1.

FIGURE 5 is a fragmentary sectional view taken on the plane indicated bythe line 5-5 in FIGURE 3.

FIGURE 6 is a cross-section of a stator blade taken on the planeindicated by the line 6--@ in FIGURE 1.

FIGURE 7 is a plot illustrating variation with altitude of corrected airflow and compressor efficiency in a typical axial-flow compressor.

FIGURE 8 is a plot illustrating typical variation of compressor losscoeflicient with Reynolds number.

FIGURE 9 is a plot illustrating the effect of the invention oncompressor efficiency.

FIGURE 10 is a cross-section of a prior art blade, either rotor orstator.

FIGURE llis a similar cross-section of a blade embodying one feature ofthe invention. 7

FIGURE 12 is an elevation view of the low pressure face of a prior artrotor blade.

FIGURE 13 is a partial sectional view of the same taken on the planeindicated by the line 1313 of FIG- URE 12.

FIGURE 14 is a view similar to FIGURE 12 illustrating the effect of asecond feature of the invention.

FIGURE 15 is a partial sectional View of the same taken on the planeindicated by the line l5--15 in FIG- URE l4.

FIGURE 1 illustrates an axial-flow compressor of known typeincorporating this invention. The compressor stator comprises an inletstructure 11, a case 12, inlet guide vanes 13, and rows of stator blades14.

The inlet structure mounts a bearing 17 supporting the forward end of arotor comprising disks 18, 19, and a tie bolt 21. Rotor blades 22 aremounted on the disks. Only the first two stages are illustrated; asmanymay be provided'as are desired.

FIGURES 2 to 5 illustrate a typical rotor blade, such as a first stageblade 22. It comprises an airfoil blade portion 26' and a dovetail root27 for mounting in disk.

13. The blade portion has a curved mean. camber line, a convex lowpressure face 28, and a concave or high pressure face 29. As will beseen from FIGURE 6,

. the stator blades 14also are airfoils with a curved mean camber, line,a convex low pressureiface 28' and a concave high pressure face 29. v "jv p The physical structure so far described is well known,

and is described more fully, in US. Patents 2,640,679 and ted StatesPatent ice 2,675,174. Such compressor structures may be improved by thepresent invention, which comprises two features, one applicable to bothrotor and stator blades, the other applicable to rotor blades only.

Before proceeding to the description of the compressor structure asmodified by the invention, reference will be made to FIGURES 10 to 15,which are generalized views illustrating the nature of the build-up ofboundary layer air on the compressor blades and the effect ofmodification of the blades according to the invention to prevent suchbuild-up.

FIGURES l0 and 11 are cross-sections of blades, which may be eitherstator or rotor blades. The blade B has a leading edge E, a trailingedge T, a low pressure face L and a high pressure face H. In the priorart blade of FIGURE 10, the spiral arrows over the rear part of the lowpressure face indicate the build-up of a thick layer of low energyviscous flow which is characteristic of operation at lower Reynoldsnumbers. FIGURE 11 illustrates a blade upon which very small ridges 34extending spanwise of the blade have been distributed chordwise of theblade on the low pressure face. The preferred structure of such anaddition to the blade will be described in detail. FIGURE 11 illustratesby the curledarrows how the spanwise ribs deflect the boundary layer airaway from the surface of the blade so that the high energy main streamair flow will scour it from the blade and carry it away. Since theboundary layer air is scoured away at several locations chordwise of theblade, it vdoes not build up into the thick layer. illustrated in FIGURE10.

So far as the rotor blades alone are concerned, there is a secondphenomenon which tends to increase the buildup of the stagnant boundarylayer. air toward the trailing edge of the low pressure face of theblade, illustrated. by FIGURES 12 and 13 showing a prior art rotorblade. The blade is identified by B and the faces and edges by the samelegends as in FIGURES l0 and 11. Theblade is shown as mounted in a rotordisk 18 in FIGURE 13. There will be a certain amount of stagnantboundary layer air adjacent the surface of the rotor 18. This will bedrawn into the low pressure area principally toward the trailing edge ofthe blade where it adds to the boundary layer air present on the face ofthe blade. Because of the rotation of the rotor there is a centrifugalforce which tends to pump this air outwardly along the blade towards thetip. However, this action does not diminishthe thicknessof the boundarylayer air, but merely moves it spanwise. The separation of the main airflow from the blade is increased by the addition of rotor boundary layerair to the blade boundary layer.

FIGURES 14 and 15 illustrate the second feature of the invention, theprovision of a rib 31 extending from the low pressure face of the bladeand preferably from the leading edge to the trailing edge as illustratedby the. arrows in these figures. The rotor boundary layer air after ashort initial travel along the surface of the blade strikes the rib 31,and is deflected outwardly into the fast: moving main air stream whichcarries it away.

The general nature of the invention being clear from the preceedinggeneral description with reference to-FIG- URES 10 to 15, we may nowproceed to the more specific description of an actual installation withreference to FIGURES 1 to 7. i

The feature of the invention applicable only to the rotor blades isprovided by a rib 31 extending substantially through the compressor. Thepreferred configuration of 3 this chrodwise-extending rib is mostclearly apparent from FIGURES 2 and 4.

The purpose of the rib is to detach boundary layer air from the lowpressure surface of the blade, particularly near the trailing edge. Asis known, a more or less stagnant boundary layer develops at the marginof the air flow path along the surface of the rotor defined by the disks18, 19, and the upper surfaces of the blade roots 27. This stagnant airrotates with the rotor. A low pressure area develops on the convex sideof the rotor blades. There also a boundary layer develops, and growstoward the trailing edge of the blade. The stagnant air at the rotorsurface tends to move outward along the face of the blade into this lowpressure area, influenced by centrifugal force, and augment the boundarylayer already present. Because of the thick boundary layer, it is notswept off by the air flowing axially through the compressor. Theaddition of the stagnant air from the inner boundary of the path,augmenting the blade boundary layer air, enhances flow separation fromthe blade and reduces efficiency.

The rib 31 intercepts this outwardly-flowing stagnant air and deflectsit into the fast-moving main air stream between adjacent blades, whichsweeps it away. Elimination of this extraneous stagnant air from thegreater part of the blade span improves the stage efficiency.

The second feature of the invention lies in means to increase theturbulence of the boundary layer air on the low pressure faces of theblades of both rotor and stator. This aspect of the invention isparticularly applicable to compressors of moderate size, rather thanlarge compressors, since the Reynolds number is less in smallcompressors.

The usual practice in compressors is to polish or otherswie finish theblades so that they have a smooth surface. This has its advantages, butit has the disadvantage that below a certain value of Reynolds number,which may be approximately 90,000, the boundary layer on the lowpressure face of the blade builds up to a considerable thickness, andthere is thus an increase in separation of flow from the low pressureface of the blade. In a typical compressor used in an aircraft engine,for example, the etficiency of the compressor decreases at higheraltitudes because of the decrease in Reynolds number. The Reynoldsnumber may vary from about 200,000 at sea level to about 50,000 at40,000 feet altitude, because of decreasing air density. The purpose ofthe present invention is to impart turbulence to the otherwise laminarboundary layer, which will minimize the growth of the boundary layer andseparation of the flow from the blade.

This result is obtained by providing a rough surface on the blade whichwill impart turbulence to the boundary layer without affecting tooadversely the air flow over the blade.

The preferred structure for this purpose comprises a number of verysmall ridges 34 extending spanwise of the blade. As will be apparentfrom FIGURES 4 and 6, these ridges are quite small compared to the spanof the blade and are of a triangular cross section. It is believed thatthe ridges may preferably be located approximately at the 7 and 3 chordvalues measured from the leading edge. As an indication of the size ofthe ridges, it is contemplated that they may be approximately $1 of aninch high.

The effect of these ridges on compressor performance may be explained byreference to FIGURES 7 to 9..

FIGURE 7 illustrates the effect of altitude on the performance of acompressor with smooth blades. The variation in corrected air flow is afunction of corrected rotor speed. However, because a turbo-prop enginegenerally operates at a nearly constant rotor speed and the airplaneMach number is fairly low, a generalized curve such as FIGURE 7 will aidin the explanation. Considering the air at sea level as standard, thecorrected air flow increases with altitude. The corrected air flow isthe weight of air flowing through the engine in unit time multiplied bythe square root of the ratio of ambient temperature to standard sealevel temperature and divided by the ratio of ambient air pressure tostandard pressure at sea level.

The efliciency of the compressor decreases very slowly with altitudefrom its sea level value up to a point at which the curve drops moresteeply with increase in altitude. If the compressor is used in a gasturbine aircraft engine, the normal flight altitude may lie beyond thedroop in the efficiency curve of FIGURE 7. One important reason for thisloss of efliciency is the decrease in Reynolds number of the air flowingthrough the compressor with altitude. This is illustrated by FIGURE 8,in which the loss coeflicient which represents loss of power in thecompressor is plotted as a function of Reynolds number. At higher valuesof Reynolds number, the loss coeflicient is nearly constant, but as theReynolds number decreases the curve begins to rise rather steeply. Thispoint, which is indicated as the break point, will occur approximatelyat a Reynolds number of 90,000. The low Reynolds numbers are thoseoccurring at high altitude and the point on the curve indicated asaltitude may represent the highest normal flight altitude of an airplaneincorporating the compressor. Gas turbine powered aircraft fly moreefliciently at high altitudes than low and, therefore, are ordinarilyflown at high altitudes such as 25,000 feet or more. This loss ofcompressor efficiency, which reduces the efiiciency of the engine and,therefore, increases its fuel consumption for a given power output, isquite undesirable.

The present invention is directed to improving the compressorefliciency. or reducing the compressor loss coeflicient, and therebyincreasing engine efliciency at high altitudes, by setting up turbulencein the boundary layer on the low pressure face of the blades. This hasan effect equivalent to an increase of Reynolds number. While theresulting disturbance of flow will decrease engine efiiciency at lowaltitude where the Reynolds number is high, it will tend to shift thebreak point to lower values of Reynolds number and thereby increase highaltitude eflicieny. This is illustrated by FIGURE 9, in which the brokenline represents the variation of efficiency of a compressor with smoothblades as a function of Reynolds number. The solid line represents thevariation of efficiency with the ridged blades of the invention. It willbe noted that a compressor incorporating the invention is somewhat lessefficient at the higher Reynolds numbers encountered at low altitude.However, it is more efficient over a considerable range of high altitudeflight in which a gas turbine aircraft normally operates. The result isthat the overall fuel consumption of the aircraft is considerablyimproved.

It will be apparent that, so far as the rotor blades are concerned, therib 31 and the ridges 34 cooperate in eliminating or preventing thedevelopment of thick, stagnant boundary layers on the low pressure faceof the rotor blades and thus cooperate to minimize separation of flowfrom the blade and promote efliciency of the compressor.

We claim:

A dynamic compressor comprising a stator and a rotor, at least onecompressor stage constituted by a row of rotor blades mounted on therotor and a row of stator blades mounted on the stator, the said bladesbeing of generally airfoil section and having a low pressure face, andridges on said blades extending spanwise on the low pressure facethereof adapted to impart turbulence to the boundary layer air thereon,the low pressure faces of the blades being smooth except as broken bythe ridges, the ridges being spaced chordwise of the blades, and atleast some of the said ridges being adjacent the mid-chord of eachblade, each rotor blade bearing a rib extending chordwise thereofadjacent to the rotor on the low pressure face thereof adapted tointercept boundary layer air flowing from the rotor spanwise of theblade and thereby prevent blanketing of the ridges by the said spanwiseflowing boundary layer air.

Clark June 28, 1932 Telfer Jan. 16, 1934 FOREIGN PATENTS Great BritainAug. 16, 1938 Great Britain Sept. 20, 1946 Great Britain June 13, 1956Great Britain Apr. 9, 1958 France July 18, 1938 France July 8, 1957

